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National Advisory Committee for Aeronautics 



NO. 90 



Research Abstracts 



SEPTEMBER 27, 1955 



CURRENT NACA REPORTS 

NACA Rept. 1197 

A STUDY OF THE CHARACTERISTICS OF HUMAN- 
PILOT CONTROL RESPONSE TO SIMULATED AIR- 
CRAFT LATERAL MOTIONS. Donald C. Cheatham. 
1954. ii, 14p. diagrs., photos., tab. (NACA 
Rept. 1197. Formerly RM L52C17) 

There are presented studies of the characteristics of 
pilot ability to control dynamically unstable yawing 
oscillations, studies of pilot control response to 
simulated aircraft yawing motions, and studies of the 
feasibility of representing pilot control response in 
an analytical form. 



NACA RM E55F28a 



STATISTICAL SURVEY OF ICING ilf^TA BittXduRED 
ON SCHEDULED AIRLINE FLIGHTS OVER THE 
UNITED STATES AND CANAD^-FJROM NOVEMBER 
1951 TO JUNE 1952. Porter J. PeKHins. 
September 1955. 44p. diagrs., RhotoSvf=2: 
(NACA RM E55F28a) ^'^*ife'^ - 

A statistical survey and a preliminary analysis are 
made in an interim report of over 600 icing en- 
counters obtained from a continuing program 
sponsored by the NACA with the cooperation of the 
airlines. Pressure-type icing-rate meters were in- 
stalled on 11 airline aircraft of various types. Icing 
conditions measured during scheduled operations 
gave relative frequencies of liquid-water content, 
icing rate, total ice accumulations, cloud tempera- 
tures, as well as horizontal and vertical extent of 
icing clouds. Liquid-water contents were higher 
than data from earlier research flights in layer -type 
clouds but slightly lower than previous data from 
cumulus clouds.' 




NACA TM 1330 

THEORY OF DYNAMIC CREEP. (K teorii 
dinamicheskoi polzuchesti). A. A. Predvoditelev 
and B. A. Smirnov. September 1955. 12p. diagr. 
(NACA TM 1330. Trans, from Moscow Universitet, 
Vestnik, v.8, no.8, 1953, p. 79-86) 

An analysis is given of the causes of the increase in 
creep under varying loads. It is suggested that the 
increase in creep is due to local rise in temperature 



over the slip planes, thus facilitating slip. A theory 
of dynamic creep is proposed, based on the Becker j, 
theory of the after-effect, which treats the metal as 
a granular structure and includes a rate factor. 
Comparison of the theory .vith experimental results 
is reserved for a future paper. ../ww 

i i 



U.9. DEPOSITORY 



NACA TN 3293 




CUMULATIVE FATIGUE DAMAGE OF AXIALLY 
LOADED ALCLAD 75S-T6 AND ALCLAD 24S-T3 
ALUMINUM-ALLOY SHEET. Ira Smith, Darnley M. 
Howard, and Frank C. Smith, National Bureau of 
Standards. September 1955. 49p. diagrs., photos., 
5 tabs. (NACA TN 3293) 

Results are presented of cumulative-fatigue-damage 
tests made on 607 specimens machined from alclad 
75S-T6 aluminum-alloy sheet 0.064 inch thick and 
198 specimens of alclad 24S-T3 and alclad 75S-T6 
aluminum-alloy sheet 0.032 inch thick. The tests of 
the 0.064-lnch-thick specimens. consistea of, 35 dif- 
ferent loading conditions and the tests oLthe 0.032- 
inch material consisted of 13 different loading -con- 
ditions. /', • /,' \> \ 
•i< // V'"\ 

NACA TN 3294 

FRICTION STUDY OF AIRCRAFT TIRE MATERIAL 
ON CONCRETE. W. G. Hample, Boeing Airplane 
Company. September 1955. 34p. diagrs., photos. 
(NACA TN 3294) 

A systematic study was made of the variation of 
frictional resistance between typical tire-tread 
material and three concrete surfaces of different 
roughness at various temperatures and normal 
pressures. The tire-tread specimens were taken 
from the thickest portion of worn ten-ply tires, and 
the three concrete test blocks were podred from the 
same mix but subjected to different surface finishes. 
Curves are presented of the apparent coefficient of 
friction as a function of normal pressure. 



NACA TN 3477 

HYDRODYNAMIC PRESSURE DISTRIBUTIONS OB- 
TAINED DURING A PLANING INVESTIGATION OF 
FIVE RELATED PRISMATIC SURFACES. Walter J. 
Kapryan and George M. Boyd, Jr. September 1955. 
82p. diagrs., photos., 5 tabs. (NACA TN 3477) 



• AVAILABLE ON LOAN ONLY. 

ADDRESS REQUESTS FOR DOCUMENTS TO NACA, 1512 H ST., NW., WASHINGTON 25, D. C, CITING CODE NUMBER ABOVE EACH TITLE; 

THE REPORT TITLE AND AUTHOR. 



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Hydrodynamic pressure distributions have been ob- 
tained during pure planing for five related prismatic 
surfaces. The distributions gave integrated lifts 
that in almost every case were well within 10 percent 
of the applied load. Comparison of experiment with 
theory shows that existing theories will adequately 
predict flat -plate pressures. For the V-shaped sur- 
faces, experiment and theory are in poor agreement. 
The lift and center-of-pressure data for both the flat 
and V-shaped surfaces are in good agreement with 
recent experimental and theoretical NACA research 
on planing surfaces. 



NACA TN 3479 

ANALYSIS OF THE HORIZONTAL -TAIL LOADS 
MEASURED IN FLIGHT ON A MULTIENGINE JET 
BOMBER. William S. Aiken, Jr. and Bernard 
Wiener. September 1955. i, 69p. diagrs., photo., 
6 tabs. (NACA TN 3479) 

Horizontal -tail loads were measured in gradual and 
abrupt longitudinal maneuvers on two configurations 
of a four-engine jet bomber. The results obtained 
have been analyzed to determine the flight values 
of the coefficients important in calculations of hori- 
zontal tail loads. The least-squares procedure used 
to determine aerodynamic tail loads from strain- 
gage measurements of structural tail loads which 
were affected by temperature is covered in detail. 
The effect of fuselage flexibility on the airplane 
motion is considered in the analysis of the abrupt - 
maneuver data. When possible, wind-tunnel results 
are compared with flight results. Some calculations 
of critical horizontal-tail loads beyond the range of 
the tests are given and compared with design loads. 



NACA TN 3486 

MEASUREMENTS OF TURBULENT SKIN FRICTION 
ON A FLAT PLATE AT TRANSONIC SPEEDS. 
Raimo J(aakko) Hakkinen, California Institute of 
Technology. September 1955. 41p. diagrs. , photo, 
tabs. (NACA TN 3486) 

The design and construction of a floating-element 
skin-friction balance are described. This instru- 
ment was applied to measurements of local skin 
friction in the turbulent boundary layer of a smooth 
flat plate at high-subsonic Mach numbers and super- 
sonic Mach numbers up to 1. 75. The principal 
difficulties which exist in comparing skin-friction 
coefficients at various Mach numbers are discussed. 



NACA TN 3491 

EXPERIMENTAL INVESTIGATION OF ECCENTRI- 
CITY RATIO, FRICTION, AND OIL FLOW OF LONG 
AND SHORT JOURNAL BEARINGS WITH LOAD- 
NUMBER CHARTS. G(eorge) B. DuBois, F(red) W. 
Ocvirk, and R. L. Wehe, Cornell University. 
September 1955. 63p. diagrs. , tabs. (NACA TN 
3491) 



NACA 
RESEARCH 



ABSTRACTS NO. 90 



The performance of plain bearings under steady 
central loading are compared and summarized by 
single-line curves covering the range of length- 
diameter ratios both above and below 1. E.xperi - 
mental date on eccentricity ratio, friction, and oil 
flow for length-diameter ratios of 1, 1-1/2, and 2 
are shovvn for comparison A'ith earlier data for 
length-diameter ratios of 1/4, 1/2, and 1. The 
combined data provide charts of plain-bearing per- 
formance which cover the practical range of length- 
diameter ratio. 



NACA TN 3493 

DEVELOPMENT OF EQUIPMENT AND OF EXPERI- 
MENTAL TECHNIQUES FOR COLUMN CREEP 
TESTS. Sharad A. Patel, Martin Bloom, Burton 
Erickson, Alexander Chwick and N(icholas) J(ohn) 
Hoff, Polytechnic Institute of Brooklyn. September 
1955. 20p. diagrs. , photos. , tab. (NACA TN 3493) 

Equipment and procedures developed for testing 
aluminum-alloy columns subjected to constant loads 
at elevated temperatures are described. Particular 
emphasis was put on determination of the influence 
of initial deviations from straightness on the critical 
time of the column, that is, the time necessary for 
the column to buckle when subjected to a constant 
load. Results are presented of tests of a number of 
2024-T4 aluminum-alloy columns having large slen- 
derness ratios. 



NACA TN 3503 

REDUCTION OF PROFILE DRAG AT SLTPERSONIC 
VELOCITIES BY THE USE OF AIRFOIL SECTIONS 
HAVING A BLUNT TRAILING EDGE. Dean R. 
Chapman. September 1955. 29p. diagrs. , photo. 
(NACA TN 3503. Supersedes RM A9H11) 

A preliminary theoretical and experimental investi- 
gation has been made on the aerodynamic character- 
istics of blunt-trailing-edge airfoils at supersonic 
velocities. The theoretical considerations indicate 
that properly designed airfoils with moderately blunt 
trailing edges can have less profile drag, greater 
lift-curve slope, and a higher maximum lift-drag 
ratio than conventional sections. These predictions 
have been substantiated by experimental measure- 
ments on airfoils of 10-percent-thickness ratio at 
Mach numbers of 1. 5 and 2. 0, and at Reynolds num- 
bers between 0. 2 and 1. 2 million. 



NACA TN 3514 

RESPONSE OF HOMOGENEOUS AND TWO- 
MATERIAL LAMINATED CYLINDERS TO SINUSOI- 
DAL ENVIRONMENTAL TEMPERATURE CHANGE, 
WITH APPLICATIONS TO HOT-WIRE ANEMOM- 
ETRY AND THERMOCOUPLE PYROMETRY. 
Herman H. Lowell and Norman (A. ) Patton. 
September 1955. ii, 143p. diagrs. , tabs. (NACA 
TN 3514) 



NACA 
RESEARCH 



ABSTRACTS NO. 90 



A theoretical investigation of the response of homo- 
geneous and tivo-material laminated, infinite cylin- 
ders to sinusoidal environmental temperature and/ 
or small heat-transfer coefficient changes was made. 
Generalized results are given for the cylinder con- 
sisting of a shell of high thermal conductivity and a 
core of loiv conductivity. The behavior of a nuniber 
of specific platinum-fused-quartz "wires" of varying 
construction and diameter exposed to a representa- 
tive airstream is indicated. For ratios of metal 
thickness to over-all radius of 0. 1, response ampli- 
tude gains of about 4. 5 are predicted as compared 
with gains of more than 10 for infinitesimal shells. 
For a relative shell thickness of 0. 05, frequency re- 
sponses of hot-wire anenion.eters, exposed-wire re- 
sistance thermometers, or thermocouples would be 
extended by at least an order of magnitude. Sinipli- 
tied analyses are included which are not exact but 
are adequate for design use. 



NACA TN 3522 

MEASUREMENTS OF THE EFFECTS OF FINITE 
SPAN ON THE PRESSURE DISTRIBUTION OVER 
DOUBLE-WEDGE WINGS AT MACH NUMBERS 
NEAR SHOCK ATTACHMENT. Walter G. Vincenti. 
September 1955. 50p. diagrs. (NACA TN 3522) 

Results are presented of measurements at low super- 
sonic speeds of the pressure distribution on two 
wings having a conimon double-wedge section and 
aspect ratios 2 and 4. Comparable results for as- 
pect ratio infinity have been published in NACA TN 
3225. The results cover the Mach number range 
from 1.166 to 1.377, which brackets the value (1.221) 
for bow-wave attachment at zero angle of attack. 
The data are discussed and compared with the previ- 
ous two-dimensional findings. 



NACA TN 3523 

THE EFFECTIVENESS OF WING VORTEX GENERA- 
TORS IN IMPROVING THE MANEIJVERING CHARAC- 
TERISTICS OF A SWEPT-WING AIRPLANE AT 
TRANSONIC SPEEDS. Norman M. McFadden, 
George A. Rathert, Jr. , and Richard S. Bray. 
Septeniber 1955. 43p. diagrs. , photos. , tab. 
(NACA TN 3523. Supersedes RM A51J18) 

The effects of wing vortex generators, multiple 
boundary-layer fences, and extension of the outer 
two segments of the wing leading-edge slats on the 
aerodynamic characteristics of a 35° swept-wing 
fighter were measured in flight tests at transonic 
speeds and high altitudes. Significant improvements 
were obtained in the pitch-up and wing-dropping- 
tendency characteristics with certain arrnagements 
of vortex generators. 



NACA TIM 3562 

VARIATION OF BOUNDARY-LAYER TRANSITION 
WITH HEAT TRANSFER ON TWO BODIES OF 
REVOLUTION AT A MACH NUMBER OF 3.12. 
John R. Jack and N. S. Diaconis. September 1955. 
16p. diagrs., photos. (NACA TN 3562) 



Cooling a cone-cylinder model to a wall-to-free- 
stream ratio of approximately 1.4 increased the 
transition Reynolds number from a value of 2.0 x 10" 
at equilibrium to 10.6 x 10". For temperature 
ratios less than 1.4, the boundary -layer flow was en- 
tirely laminar. For a parabolic -nosed body, the 
transition Reynolds number was about twice that of 
the cone-cylinder model over the temperature range 
investigated. 



NACA TN 3563 

HEAT LOSS FROM YAWED HOT WIRES AT SUB- 
SONIC MACH NUMBERS. Virgil A. Sandborn and 
James C. Laurence. September 1955. 44p. 
diagrs., photo. (NACA TN 3563) 

Heat-loss data at angles of yaw and fixed subsonic 
Mach numbers for several wires of different diam- 
eters commonly used in hot-A'ire anemometry are 
presented. Possible methods of correlating the 
data are examined. The relation of the Reynolds 
number normal to the flow, .vhich has been used by 
rriost researchers, was inadequate except near a 
Mach number of zero. An empirical relation based 
on weighted addition of the heat losses of wires 
normal and parallel to the flow correlated all data 
reasonably well. 



NACA TN 3566 

A POLAR-COORDINATE SURVEY METHOD FOR 
DETERMINING JET-ENGINE COMBUSTION- 
CHAMBER PERFORMANCE. Robert Friedman and 
Edward R. Carlson. September 1955. 29p. 
diagrs., photo., tab. (NACA TN 3566) 

An automatic polar-coordinate traversing system is 
described that sweeps a probe through a quarter- 
annular exhaust duct circumferentially at selected 
radial positions. With a single combined pressure 
and temperature probe, temperature and pressure 
are recorded simultaneously as a function of probe 
position. The use of these data in calculating 
temperature and flow profiles, combustion efficiency, 
and pressure loss is shown. 



BRITISH REPORTS 



N-38605* 

Aeronautical Research Council (Gt. Brit. ) 
THE USE OF QUARTZ IN THE MANUFACTURE OF 
SMALL DIAMETER PITOT TUBES. J. R. Cooke. 
1955. 14p. diagrs., photos., tab. (ARC CP 193) 



This note describes the method of manufacture of 
small quartz-tipped pitot tubes (doA-n to 0.005 in. 
outside tip diameter) which have been used success- 
fully for boundary-layer measurements on small 
models in a supersonic wind tunnel. Tests have 
been made of the effects of taper and end finish on 
the accuracy of measurement, and of the effect of 
the inside diameter of the tip (for a standard taper) 
on response rate. For a given inside tip diameter 
the tapered quartz tubes gave a faster response rate 
than the stainless steel hj-podermic tubes previously 
used. 



N-38606* 

Aeronautical Research Council (Gt. Brit. ) 
A NOTE ON THE SOUND FROM WEAK DISTURB- 
ANCES OF A NORMAL SHOCK WAVE. Alan Powell 
1955. lOp. diagrs. (ARC CP 194) 

The disturbances of a shock wave by sound waves or 
temperature fluctuations are studied in one dimen- 
sion to a first-order approximation. In general, 
both sound waves and temperature fluctuations arise 
behind the shock ivave. Expressions are given for 
their amplitudes and calculated for > = 1.4. Sound 
waves colliding with the shock wave are amplified, 
but sound waves are almost annihilated by weak 
shock waves if originally travelling in the same di- 
rection as the shock wave. Small temperature 
fluctuations give rise to much sound on an acoustical 
scale. 



N-38607* 

Aeronautical Research Council (Gt. Brit.) 
REQUIREMENTS FOR UNIFORMITY OF FLOW IN 
SUPERSONIC WIND TUNNELS. D. E. Morris and 
K. G. Winter. 1955. 9p. diagr. (ARC CP 197) 

An analysis is made of the effects of nonuniformity 
of flow on the pressure measurements on the surface 
of a model and also on the force and moment meas- 
urements. The following standards of flow uniform- 
ity are derived - variations in flow direction to be 
less than ±0.1° in the range M = 1.4 to 3; variation 
in Mach number to be less than ±0.003 at M = 1.4 
increasing to +0.01 at M = 3. A brief analysis is 
made of the errors in model manufacture and their 
effects on force and pressure measurements. Using 
the same standards as were used in deducing the 
requirements for flow uniformity quoted above, it is 
concluded that present standards of model manu- 
facture are satisfactory overall, though for accurate 
pressure plotting tests at lo.v supersonic Mach num- 
bers a higher standard is desirable. 



N-38608* 

Aeronautical Research Council (Gt. Brit.) 
A CRITERION FOR THE PREDICTION OF THE RE- 
COVERY CHARACTERISTICS OF SPINNING AIR- 
CRAFT. T. H. Kerr. 1955. 22p. diagrs., tabs. 
(ARC CP 195) 



NACA 

RESEARCH ABSTRACTS NO. 90 



It has been deduced that the t.vo most in.portant pa- 
rameters are the unbalanced roUing-n.on.ent coeffi- 
cient about the wind a.xis in the spin and the ratio of 
pitching to rolling moment of inertia. Using the 
results of full-scale spinning tests on 33 aircraft, it 
has been possible to establish empirical relation- 
ships between the estimated unbalanced rolling- 
moment coefficient and the inertia ratio which effect- 
ively divide the aircraft into the three groups which 
have satisfactory, borderline, and unsatisfactory re- 
covery characteristics. A simple method is pre- 
sented for estimating the unbalanced rolling-moment 
coefficient knowing only the shape of the aircraft. 
The euipirical relationships should give a good indi- 
cation of the spin-recovery characteristics on new 
designs. 

N-38616* 

Aeronautical Research Council (Gt. Brit.) 
MODEL TESTS ON THE EFFECTS OF SLIPSTREAM 
ON THE FLOW AT VARIOUS TAILPLANE POSI - 
TIONS ON A FOUR-ENGINED AIRCRAFT. PART L 
TESTS WITH CONTRA -ROTATING PROPELLERS. 
D. E. Hartley, A. Spence. and D. A. Kirby. 
PART n. TESTS WITH SINGLE ROTATING PRO- 
PELLERS. D. A. Kirby. 1955. 37p. diagrs., 
tabs. (ARC R & M 2747; ARC 12,355; ARC 14, 166. 
Supersedes RAE Aero 2322; RAE Aero 2322a) 

Systematic wind-tunnel tests have been made to in- 
vestigate the effects of slipstream on the flow near 
the tail plane of a typical civil transport with four 
contra-rotating propellers. Tail-plane height has 
been varied for each of several wing-body arrange- 
ments; only one tail plane and one propeller position 
have been used. This report presents the main re- 
sults in the form of changes in mean downwash angl^ 
and velocity at the tail plane, as functions of tail- 
plane position, lift coefficient, and propeller thrust. 



N-38617* 

Aeronautical Research Council (Gt. Brit.) 
DETERMINATION OF THE STRESS DISTRIBUTION 
IN REINFORCED MONOCOQUE STRUCTURES. 
PART I A THEORY OF FLAT-SIDED STRUC- 
TURES. L. S. D. Morley. 1955. 23p. diagrs., 
photos. (ARC R &M 2879; ARC 14.814. Superse- 
des RAE Structures 120) 

This paper is concerned with the estimation of the 
stress distribution in the neighborhood of a discon- 
tinuity in reinforced monocoque flat-sided struc- 
tures. A theory is given based upon a shell model 
possessing uniformly distributed stringers but dis- 
crete ribs, which can serve as a basis for the prac- 
tical solution of a wide range of flat-sided struc- 
tures such as rectangular or polygonal fuselages and 
wing boxes. 



N-38618* 

Aeronautical Research Council (Gt. Brit. ) 
THE THEORETICAL WAVE DRAG OF SOME 
BODIES OF REVOLUTION. L. E. Fraenkel. 1955. 
26p. diagrs., tab. (ARC R & M 2842; ARC 14, 334. 
Supersedes RAE Aero 2420) 



NACA 

RESEARCH ABSTRACTS NO. 90 



This report investigates the wave drag of bodies of 
revolution with pointed or open-nose forebodies and 
pointed or truncated afterbodies. The "quasi- 
cylinder" and "slender-body" theories are reviewed 
a reversibility theorem is established, and the con- 
cept of the interference effect of a forebody on an 
afterbody is introduced. The theories are applied 
to bodies whose profiles are either straight or para- 
bolic arcs, formulas and curves being given for 
forebody and afterbody drag, and for the interfer- 
ence drag. The results of the two theories are com- 
pared and are seen to agree well in the region of 
geometries where both theories are applicable. 



N-38619* 

Aeronautical Research Council (Gt. Brit. ) 
AN EXPERIMENTAL INVESTIGATION OF STRESS 
DIFFUSION IN NON-BUCKUNG PLATES. L. H. 
Mitchell. 1955. 20p. diagrs., photos. (ARC 
R & M 2878. Supersedes ARC 14, 934; Strut 1540) 

This report provides experimental results for com- 
parison with theoretical analyses of stress diffusion 
problems. The structures considered consist of 
plane reinforced sheet which has been assumed not 
to buckle. Symmetrical loads are applied to the 
edge booms connected to the sheet by continuous no- 
slip joints. Attention is concentrated on the stress 
distribution near the ends of the parallel strips of 
plate. An outline of the existing theoretical work 
which is applicable to this type of problem is given. 
The stringer-sheet theory is compared with the 
photoelastic results. Some attention is also given 
to transverse end stiffeners which seem to have 
little effect on the shear stresses. 



N-38620* 

Aeronautical Research Council (Gt. Brit. ) 
THE BOUNDARY LAYER WITH DISTRIBUTED SUC- 
TION. M. R. Head. 1955. lOOp. diagrs., photos, 
tabs. (ARC R & M 2783. Supersedes ARC 13, 897; 
FM 1547; Perf. 771) 

Experiments performed in flight at Reynolds num- 
bers in the region of 3 x 10° have clearly demon- 
strated the stabilizing effect of small amounts of 
distributed suction on the laminar boundary layer. 
In the absence of a pressure gradient and in adverse 
gradients similar to those occurring on a normal 
airfoil, transition of the bouiidary layer to the tur- 
bulent form has been prevented by the use of such 
suction quantities as may be expected to lead to very 
considerable reductions in effective drag. It ap- 
pears, however, that for extensive laminar flow to 
be achieved in this way, the surface must be free 
from such excrescences as would cause transition 
in the absence of suction. 



N-38621* 

Aeronautical Research Council (Gt. Brit. ) 
METHODS FOR CALCULATING THE LIFT DISTRI- 
BUTION OF WINGS (SUBSONIC LIFTING-SURFACE 
THEORY). H. Multhopp. 1955. 96p. diagrs., 
tabs. (ARC R & M 2884; ARC 13, 439. Supersedes 
RAE Aero 2353) 

These methods for calculating the load distribution 
on wings of any plan form are based on the concep- 
tions of lifting-surface theory. Computer work 
time is shortened by careful choice of the positions 
of pivotal points, by plotting once for all those parts 
of the downwash integral which occur frequently and 
by a consequent application of approximate integra- 
tion methods similar to those devised by the author 
for lifting-line problems. The basis of the method 
is to calculate the local lift and pitching moment at 
a number of chordwise sections from a set of linear 
equations satisfying the downwash conditions at two 
pivotal points in each section. 



N-3871f 

Aeronautical Research Council (Gt. Brit.) 
SIMPLE EVALUATION OF THE THEORETICAL 
LIFT SLOPE AND AERODYNAMIC CENTRE OF 
SYMMETRICAL AEROFOILS. H. C. Garner. 
1955. 20p. tabs. (ARC R & M 2847. Supersedes 
ARC 14,337; Perf. 847; S & C 2561) 

This paper presents a simple method of calculating 
theoretical values of the lift slope (a^)™ and the 
position of aerodynamic center hf in two- 
dimensional incompressible flow. Starting with the 
ordinates of an airfoil, the method in section 3 pro- 
vides first and second approximations to both de- 
rivatives, which are compared with exact theory 
and other calculated values in Tables 2 and 3 for 
various symmetrical airfoils listed in Table 1. In 
section 5 a correction to the first approximation is 
introduced so as to permit the evaluation of (aj^)™, 

within 1/2 percent and hf within about 0.001 in 
less than a quarter of an hour. A complete illus- 
trative calculation is set out in Table 4. 



N-38712* 

Aeronautical Research Council (Gt. Brit. ) 
BOUNDARY-LAYER CONTROL FOR HIGH LIFT BY 
SUCTION AT THE LEADING-EDGE OF A 40 DEG 
SWEPT-BACK WING. E. D. Poppleton. 1955. 
38p. diagrs., tabs. (ARC R &M 2897; ARC 14,771. 
Supersedes RAE Aero 2440) 

Wind-tunnel tests on the 10-percent-thick, constant - 
chord, aspect-ratio-4.6 wing are discussed. 
Boundary-layer control was applied along the whole 
leading edge; a comparison was made between the 
effects of distributed suction and suction through a 



slot. A 45-percent Fowler flap was used in some 

tests. The overall effect of the two systems was 

similar, giving an increase in Cr by increas- 

'-'max 

ing the stalling angle of attack and making the wing 
statically stable up to the stall, when there was a 
severe loss of lift. The tests were designed to de- 
termine whether leading-edge suction would produce 
comparable increases in C^ on swept wings 

and, also whether tip stall could be prevented. 



N-38713* 

Aeronautical Research Council (Gt. Brit. ) 
ON THE APPLICATION OF OBLIQUE CO- 
ORDINATES TO PROBLEMS OF PLANE ELAS- 
TICITY AND SWEPT-BACK WING STRUCTURES. 
W. S. Hemp. WITH AN APPENDIX. S. R. Lewis. 
1955. 46p. diagrs., tabs. (ARC R & M 2754; ARC 
12, 981. Supersedes College of Aeronautics Rept. 
31; College of Aeronautics Rept. 44) 

Methods are discussed by which designers can solve 
problems of stress distribution and deflection for 
the case of sweptback wing structures whose ribs 
lie parallel to the direction of flight. The mathe- 
matical basis is developed and formulas are derived. 
The results are applied to a uniform, symmetrical, 
rectangular section sweptback box. Theories of 
stress distribution and deflections are obtained for 
the case of loading by normal forces and couples 
applied to the ends of the box. The main results 
are then generalized to cover the case of a more 
representative wing structure. Functions useful in 
the application of the theory are given in an appendix. 



N-38714* 

Aeronautical Research Council (Gt. Brit. ) 
LOW-SPEED TUNNEL MODEL TESTS ON TAIL- 
PLANE ROLLING MOMENTS IN SIDESLIP. 
A Spence, J. W. Leathers, and D. A. Kirby. 1955. 
20p. diagrs., tabs. (ARC R& M 2941; ARC 14,701. 
Supersedes RAE Tech. Note Aero 2123) 

Measurements were made of the effect of sideslip on 
the rolling moment on a 41.5° sweptback tail plane 
mounted at three heights on the fin of a model of a 
single jet aircraft with a 40° sweptback wing. Inci- 
dence and tail-plane setting were varied, and the ef- 
fects of rudder deflection were obtained with the 
tail plane at the top of the fin. Brief results on a 
delta aircraft model with a delta tail plane at the top 
of the fin are also included. Values of the rolling 
moment on the tail plane were obtained from meas- 
urements of the bending moment on the starboard 
half of the tail plane about a hinge just outside the 
fin. 



NACA 

RESEARCH ABSTRACTS NO. 90 



N-38715* 

Aeronautical Research Council (Gt. Brit.) 
TWO-DIMENSIONAL CONTROL CHARACTER- 
ISTICS. L. W. Bryant, A. S. Halliday, and A. S. 
Batson. 1955. 47p. diagrs. (ARC R & M 2730. 
Supersedes ARC 13,039; S & C 2385; ARC 13,065; 
S & C 2386) 

Researches on the lift, pitching moments, and hinge 
moments of airfoils with plain flaps have been car- 
ried out at the National Physical Laboratory at a 
Reynolds number of about 10°. The results have 
been presented in a generalized form, which shows 
promise of being applicable over a wide field. It 
appears that a suggestion due to Preston that the 
ratio of experimental lift slope (dCL/da = ai) to the 
theoretical value (ai)-T., corresponding to the 
Joukowsky condition of flow past the trailing edge, 
provides a criterion giving the combined effects of 
Reynolds number, transition points, and airfoil 
shape on dCi^/da, and is a very useful starting 
point for the estimation of control characteristics. 



N-38716* 

Aeronautical Research Council (Gt. Brit. ) 
PERMISSIBLE DESIGN VALUES AND VARIABILITY 
TEST FACTORS. R. J. Atkinson. 1955. 20p. 
diagrs., tabs. (ARC R & M 2877; ARC 11,619; 
ARC 13,748. Supersedes RAE Tech. Note Structures 
15; RAE Tech. Note Structures 61) 

For the design of structural elements it is postulated 
that: not more than 10 percent of any given design 
should have strength below the design value, and not 
more than 0.1 percent should have strength below 90 
percent of the design value. This rule forms a 
working basis for the interpretation of tests on sta- 
tistical lines. On the basis of a fixed probability the 
report deduces: expressions for the derivation of 
permissible design values from a given number of 
test results, the number of test results required so 
that the estimates of permissible design values can 
be regarded as sufficiently accurate, and the factor 
which should be applied to the results of tests on any 
number of similar components designed to meet a 
specified requirement. 



N-38717* 

Aeronautical Research Council (Gt. Brit.) 
IMPROVEMENTS IN THE FATIGUE STRENGTH OF 
JOINTS BY THE USE OF INTERFERENCE FITS. 
W. A. P. Fisher and W. J. Winkworth. 1955. 17p. 
diagrs., photos., tabs. (ARC R & M 2874; ARC 
15,014. Supersedes RAE Structures 127) 



NACA 

RESEARCH ABSTRACTS NO. 90 



Fatigue test results are given for aluminum alloy 
flat bars with a single hole loaded by a pin in double 
shear. In one series the pin was fitted directly in 
the hole with various degrees of interference fit up 
to 0.003 in. excess diameter. The other series had 
a mild steel bush interposed with similar degrees of 
interference in the bar, but with a push fit between 
pin and bush. Both sets showed a great increase in 
fatigue strength for interference fits above a critical 
value. 



N-38718* 

Aeronautical Research Council (Gt. Brit. ) 
AN EXAMINATION OF THE FLOW AND PRESSURE 
LOSSES IN BLADE ROWS OF AXIAL-FLOW TUR- 
BINES. D. G. Ainley and G. C. R. Mathieson. 
1955. 33p. diagrs. (ARC R & M 2891; ARC 
14,232. Supersedes NOTE R. 86) 

Available information is studied and analyzed to de- 
termine magnitudes of gas pressure losses and de- 
flections in a wide variety of blade rows and to de- 
termine the separate influences of variables such as 
blade shape, blade spacing, gas Mach number, 
Reynolds number, incidence, etc, Special attention 
is paid to "secondary losses. ' Effects of blade tip 
clearance are also considered. Empirical guiding 
rules and charts are derived from which approximate 
values of the overall pressure losses and gas deflec- 
tions in a range of blade rovvs can be deduced. It is 
found that secondary losses can in many instances be 
large, the loss being generally found to be great 
when the blading has low reaction. 



N-38719* 

Aeronautical Research Council (Gt. Brit. ) 
FLUTTER AND RESONANCE CHARACTERISTICS 
OF A MODEL CANTILEVER WING CARRYING 
LOCALISED MASSES. N. C. Lambourne. 1955. 
25p. diagrs., tabs. (ARC R & M 2866. Supersedes: 
ARC 13,910; 0.939; ARC 11,008; 0.687) 

Resonance tests on a model cantilever wing carrying 
concentrated masses were made in conjunction with 
flutter tests. Measurements were made with 
masses up to approximately five times the mass of 
the bare wing added at two positions. Flutter and 
resonance characteristics are placed in juxtaposi- 
tion. An attempt is made to correlate the two sets 
of phenomena by means of the Kiissner criterion. 
Distortion modes of flutter are analyzed into normal 
mode components. Results suggest that for a wing 
rigidly fixed at the root and carrying a single con- 
centrated mass the first three normal modes are 
sufficient to define the flutter mode. 
Copies obtainable from NACA, Washington 



N-38720* 

Aeronautical Research Council (Gt. Brit. ) 
SOME APPLICATIONS OF THE LAME FUNCTION 
SOLUTIONS OF THE LINEARISED SUPERSONIC 
FLOW EQUATIONS. PART I - FINITE SWEPT- 
BACK WINGS WITH SYMMETRICAL SECTIONS AND 
ROUNDED LEADING EDGES. PART E - CAMBER- 
ED AND TWISTED WINGS. G. M. Roper. 1955. 
42p. diagrs. (ARC R & M 2865; ARC 14,473; ARC 
14,475; ARC 14,476. Supersedes RAE Aero 2436; 
RAE Aero 2437) 

In the present paper some special solutions are 
found. Some of these solutions are combined with 
previous solutions to give (a) pressure distribution 
and wave drag at zero lift on some finite unyawed 
sweptback wings having symmetrical sections with 
rounded leading edges and wing tips perpendicular to 
the wind direction, and (b) the change in pressure 
distribution and wave drag at zero lift on the surface 
of a Squire wing when the thickness chord ratio is 
modified. Some additional solutions applicable to 
cambered and twisted wings are also given. 



N-3872f 

Aeronautical Research Council (Gt. Brit.) 
THE APPLICATION OF THE EXACT METHOD OF 
AEROFOIL DESIGN. M. B. Glauert. 1955. 45p. 
diagrs., tabs. (ARC R&M 2683. Supersedes 
ARC 10,933; FM 1161) 

This report considers in detail the design of air- 
foils by Lighthill's exact method, in which the ve- 
locity over the airfoil surface is prescribed as a 
function of the angular coordinate on the circle into 
which the airfoil may be transformed. The mathe- 
matical basis of the method is set out, means for 
obtaining desired characteristics for the airfoil are 
developed, and the procedure to be followed in the 
actual design is fully discussed. Various special 
functions are introduced to increase the range and 
practical utility of the velocity distributions obtain- 
able, and these and other functions are fully tabu- 
lated. The calculations for the design of a particu- 
lar thick suction airfoil are set out in detail. 
Copies obtainable from NACA, Wasliington 



N-38722* 

Aeronautical Research Council (Gt. Brit. ) 
AN EXPERIMENTAL INVESTIGATION OF THE 
BOUNDARY LAYER ON A POROUS CIRCULAR 
CYLINDER. D. G. Hurley and B(rian) Thwaites. 
1955. 14p. diagrs., photos. (ARC R&M 2829. 
Supersedes ARC 14,158; FM 1584) 



The report describes an experimental investigation 
of the boundary layer on the surface of a porous 
circular cylinder at which there is a normal inward 
velocity. The primary object of the experiments 
was to test the approximate theory of reference 1 for 
calculating the development of a laminar boundary 
layer under conditions of continuous suction. The 
formula given in that reference for calculating the 
momentum thickness of the layer gave results in ac- 
cord with the experimental determinations. Owing 
to practical difficulties in the exploration of the very 
thin boundary layers and in the determination of the 
velocity gradient around the surface, other com- 
parisons with the theory were difficult. 



N-38723* 

Aeronautical Research Council (Gt. Brit. ) 
FORMULAE FOR ESTIMATING THE FORCES IN 
SEAPLANE-WATER IMPACTS WITHOUT ROTA- 
TION OR CHINE IMMERSION. R. J. Monaghan and 
P. R. Crewe. 1955. 28p. diagrs., tabs. (ARC 
R & M 2804; ARC 12,399. Supersedes RAE 
Aero 2308) 

This report contains design formulas for estimating 
the maximum forces, together with the times and 
drafts associated with these forces, in main-step 
landings of seaplanes provided there is neither ro- 
tation nor chine immersion. Good agreement is 
formed with the results of model tests made under 
controlled conditions at NACA. The basic formulas 
and curves presented are considered to be the most 
satisfactory and accurate of the many proposed in 
recent years. They involve the use of a new basic 
parameter which is a measure of the effect of 
forward velocity; a new formula for associated 
mass, and a new method of plotting which is con- 
sidered to be the most useful for the analysis of ex- 
perimental data. 



N-38724* 

Aeronautical Research Council (Gt. Brit. ) 
WIND-TUNNEL TESTS ON THE NACA 63A009 
AEROFOIL WTTH DISTRIBUTED SUCTION OVER 
THE NOSE. N. Gregory and W. S. Walker. 1955. 
17p. diagrs., tabs. (ARC R & M 2900. Supersedes 
ARC 15,184; Perf. 987; FM 1787) 

The effects of distributed suction on the stalling 
characteristics of the airfoil are described. The 
most economical extent of suction was from the lead- 
ing edge for 2. 75 percent chord round the upper 
surface. At a R = 1.15 x 10^, a suction-quantity 
coefficient of 0.0034 increased CLmax from 0.86 
to 1.50 by delaying the stall from a = 11° to 
a = 20°. Scale effect on the flow was investigated 
at a = 14°. The airfoil was also tested with a 
20-percent split flap at 60° deflection. Suction 
gave half the increase on the flapped airfoil that it 
gave on the plain airfoil. The airfoil was modified 
for further testing by reducing the chord and 
blunting the nose. 



NACA 
RESEARCH 



ABSTRACTS NO. 90 



N-38725* 

Aeronautical Research Council (Gt. Brit.) 
DETAILED OBSERVATIONS MADE AT HIGH IN- 
CIDENCES AND AT HIGH-SUBSONIC MACH NUM- 
BERS ON GOLDSTEIN 1442/1547 AEROFOIL. 
H. H. Pearcey and M. E. Faber. 1954. 52p. 
diagrs., photos., tabs. (ARC R & M 2849. Super- 
sedes ARC 13,531; FM 1498; Perf. 714) 

Surface-pressure distribution, shock-wave photo- 
graphs, and observations of boundary-layer separa- 
tion have been made over a wide range of angle of 
attack. The observations enable the effects of 
compressibility on CLmax ^"^ °" "^^ nature of the 
stall to be studied in detail for the two-dimensional 
case. The pitching-moment coefficients, also, can 
be integrated from the pressure distributions. Cer- 
tain features of the results are thought to be of fairly 
general interest and application. 



N-38728* 

Royal Aircraft Establishment (Gt. Brit.) 
TECHNIQUES FOR THE MEASUREMENT OF THE 
AERODYNAMIC FORCES ON OSCILLATING AERO- 
FOILS. W. G. Molyneux. June 1955. 30p. diagrs. 
(RAE Tech. Note Structures 161) 

The various techniques for oscillatory force meas- 
urements are considered in relation to their applica- 
tion to the measurement of the aerodynamic coeffi- 
cients for a rectangular wing oscillating in modes of 
vertical translation and uniform pitch. It is shown 
that the eight relevant coefficients L^, L^, L^, 
L^, M^, M^, M(j and M^ are obtainable by any 
of the techniques described. The survey is not ex- 
haustive, but it provides a basis for comparison of 
the various techniques and should be of assistance to 
investigators in this field in indicating the particular 
technique most likely to meet their requirements. 



N-38729* 

Royal Aircraft Establishment (Gt. Brit.) 
THE EFFECT OF WATER ON THE POROSITY OF 
PARACHXrrE FABRICS. J. E. Swallow. May 
1955. 18p. diagrs., tabs. (RAE Tech. Note 
Chem. 1248) 

Air flow through parachute fabrics was found to be 
seriously affected by water. The porosity of the 
nylon, cotton, Fortisan and Terylene fabrics ex- 
amined was decreased and became negligible for the 
closer weaves. This was mainly a surface tension 
effect, but swelling was a contributory factor for 
cellulosic fabrics. Mock-leno weave nylon fabrics 
were least affected. 



NACA 
RESEARCH 



N-38730* 



ABSTRACTS NO. 90 



Royal Aircraft Establishment (Gt. Brit. ) 
ON THE INTEGRAL EQUATIONS OF TWO DIMEN- 
SIONAL SUBSONIC FLUTTER DERIVATIVE 
THEORY. D. E. Williams. June 1955. 39p. 
(RAE Structures 181) 

This note gives the result of an attempt to find an 
analytical solution of Possio's integral equation - 
the equation \vhich connects the downwash and the 
pressure distribution on an airfoil oscillating in two- 
dimensional subsonic compressible flow. A method 
is given for solving this problem and for solving the 
corresponding problem in incompressible floiv - the 
solution of Birnbaum's integral equation. 



N-38732* 

Royal Aircraft Establishment (Gt. Brit. ) 
THE DETERMINATION OF FLUORINE IN ORGANIC 
COMPOUNDS CONTAINING FLUORINE AND PHOS- 
PHORUS. T. R. F. W. Fennell. May 1955. lip. 
diagr., tabs. (RAE Tech. Note Chem. 1251) 

A published method for the determination of fluoride 
in the presence of phosphate ion has been found to 
yield erroneous results. The method has been 
modified to overcome this fault. 



N-38759* 

Aeroplane and Armament Experimental Establish- 
ment (Gt. Brit. ) THE EFFECT OF THE GROUND 
ON A HELICOPTER ROTOR IN FORWARD FLIGHT. 
I. C. Cheeseman and W. E. Bennett. July 11, 1955. 
13p. diagrs. (AAEE/Bes/288) 

An approximate method of estimating the effect of 
the ground on the lift of a rotor at any forward speed 
is described. Flight tests on several different air- 
craft show reasonable agreement with the theory. 
Curves are given showing the relation between 
thrust, height, speed, and power. The theory has 
been extended to include the effect of a variation in 
blade loading and shows that within the range that 
this parameter takes on present single rotor heli- 
copters the effect is small. 



N-38761* 

Royal Aircraft Establishment (Gt. Brit. ) 
A UNIFIED THEORY OF PERFECTLY PLASTIC 
PLATES. E. H. Mansfield. May 1955. 53p. 
diagrs. (RAE Structures 170) 

A theory is developed for determining the collapse 
load and the collapse mechanism for perfectly plas- 
tic plates under normal loading. A number of solu- 
tions to simple problems is first presented and the 
theory is extended to deal with plates of arbitrary 
plan carrying a concentrated load, and to plates of 
rectangular or regular polygonal plan carrying a 
uniformly distributed load. 



N-38781* 

Forest Products Research Lab. (Gt. Brit.) 
INVESTIGATIONS INTO GLUES AND GLUING. 
PROGRESS REPORT EIGHTY-FIVE - JUNE 1955. 
BEHAVIOR OF GLUED WOOD PRODUCTS IN LIGHT 
NAVAL CRAFT. PART I - SYNOPTIC REPORT. 
FIFTH YEAR'S ANALYSIS. R. J. Newall and L. S. 
Doman. 6p. (Forest Products Research Lab. 
Supersedes corresponding part of Progress Report 
71) 

This investigation consists in storing samples of ply- 
wood and other glued wood products in selected loca- 
tions for periods up to 10 years. At intervals, 
samples are removed and systematically tested for 
deterioration of the glue lines, fungal attacK, etc. 
Inspections have been made at six-monthly intervals 
over the past 5 years and a summary of the observa- 
tions is presented. 



N-38782* 

Forest Products Research Lab. (Gt. Brit.) 
COMPOSITE WOOD SECTION. TRIALS OF 
TIMBERS FOR PLYWOOD MANUFACTURE. 
ANINGUERLA - ANINGUERIS ALTISSIMA - UGANDA. 
(NO RELIABLE WEIGHT FIGURES AVAILABLE BUT 
PROBABLY BETWEEN 35 AND 40 LB. PER CUBIC 
FOOT AT 15 PER CENT MOISTURE CONTENT). 
(PROGRESS REPORT TWENTY-EIGHT). June 1955. 
12p. tabs. (Forest Products Research Lab. ) 



N-38783* 

Forest Products Research Lab. (Gt. Brit.) 
COMPOSITE WOOD SECTION. TRIALS OF 
TIMBERS FOR PLYWOOD MANUFACTURE. ABURA 
(NZINGU)-MITRAGYNA STIPULOSA - UGANDA. 
(36 POUNDS PER CUBIC FOOT AT 15 PER CENT 
MOISTURE CONTENT). (PROGRESS REPORT 
TWENTY-SEVEN). June 1955. 14p. tabs. (Forest 
Products Research Lab.) 



N-38784* 

Forest Products Research Lab. (Gt. Brit. ) 
COMPOSITE WOOD SECTION. TRIALS OF 
TIMBERS FOR PLYWOOD MANUFACTURE. 
DAHOMA - PIPTADENIA AFRICANA - UGANDA. 
(47 POUNDS PER CUBIC FOOT AT 15 PER CENT 
CONTENT). MUCHENCHE - PIPTADENIA 
BUCHANANH - UGANDA. (35 POUNDS PER CUBIC 
FOOT AT 15 PER CENT MOISTURE CONTENT). 
(PROGRESS REPORT TWENTY-SDC). June 1955. 
lip. tabs. (Forest Products Research Lab. ) 



N-38785* 

Forest Products Research Lab. (Gt. Brit.) 
MOISTURE RELATIONS OF COMPOSITE WOOD 
PRODUCTS. PROGRESS REPORT TWENTY- 
SEVEN - JUNE 1955. THE FURROWING OF 
VENEERED BLOCKBOARD. J. F. S. Carruthers. 
9p. diagrs., tabs. (Forest Products Research Lab. 
Supersedes Progress Report 26, May, 1954) 



10 




3 1262 08153 278 9 



NACA 
RESEARCH 



ABSTRACTS NO. 90 



The purpose of this investigation was to determine 
the cause of the furrowing which sometimes occurs 
on the surface of veneered blockboard after polishing. 
Three different core constructions were employed 
and an explanation of the furrovving is given for each 
type. 



N-38807* 

Royal Aircraft Establishment (Gt. Brit. ) 
VELOCITY CALCULATIONS BY CONFORMAL MAP- 
PING FOR TWO-DIMENSIONAL AEROFOILS. 
D. A. Spence and N. A. Routledge. February 1955. 
48p. diagrs., tabs. (RAE Aero 2539) 

A method is derived for computing the conformal 
transformation between the plane of an airfoil of 
arbitrary shape (symmetrical or cambered), and the 
plane of its velocity potential at zero lift (in which 
the airfoil contour becomes a slit), in order to per- 
mit calculations of the velocity at points off the sur- 
face. The integral equation which relates the con- 
tours is derived by an application of Cauchy's 
theorem, and solved to the order of the square of 
thickness ratio. The solution is found by repre- 
senting the ordinate distribution by a Fourier series. 
The rapid tailing-off of the Fourier coefficients for 
all smooth airfoil shapes then leads to high accuracy 
being achieved with a comparatively small amount of 
effort. The method is straightforward and has 
proved easy to use. 



N-38808* 

Royal Aircraft Establishment (Gt. Brit. ) 
THE CHARACTERISTIC FREQUENCIES OF SMALL 
OSCILLATIONS IN THE FLOW PAST BLUFF 
BODIES. D. A. Spence. May 1955. 23p. diagrs. 
(RAE Aero 2532) 

Summary: When a bluff body is placed in a steady 
stream it experiences buffeting, the periodicity of 
which can be explained in terms of interactions 
between external and boundary layer regions. It 
is shown that the frequency must satisfy a character- 
istic equation in order for the oscillations induced in 
the boundary layer to be compatible with those in the 
outside stream. The equation is derived formally 
for Lighthill's step case and for that of the circular 
cylinder. The Karman vortices which are observed 
in the latter case appear to be a consequence of the 
oscillatory character of the circulation around the 
cylinder. 



DECLASSIFIED NACA REPORTS 



NACA RM A54F28 

ON THE RANGE OF APPLICABILITY OF THE 
TRANSONIC AREA RULE. John R. Spreiter. 
August 1954. 21p. (NACA RM A54F28) (Declassi- 
fied from Confidential, 9/7/55) 



Some insight into the range of applicability of the 
transonic area rule has been gained by comparison 
with the appropriate similarity rule of transonic 
flow theory and with experimental data for a large 
family of rectangular wings having NACA 63AXXX 
profiles. 



NACA RM A54J07 

THEORETICAL PRESSURE DISTRIBUTIONS FOR 
SOME SLENDER WING-BODY COMBINATIONS AT 
ZERO LIFT. Paul F. Byrd. January 1955. 39p. 
diagrs. (NACA RM A54J07) (Declassified from 
Confidential, 9/7/55) 

Theoretical calculations are made of the pressure 
distributions for some slender, symmetrical wing- 
body combinations in subsonic and sxipersonic flow. 
The combinations consist first of nonlifting, swept - 
t)ack wings mounted on a circular cylinder and 
second of such wings mounted on a body indented so 
that the local cross-sectional area of the combina- 
tion is constant. The results indicate that indenta- 
tion straightens out the isobars along the wing and 
diminishes the maximum perturbation velocities. 



NACA RM L52H08 

A STUDY OF THE ZERO-LIFT DRAG-RISE CHAR- 
ACTERISTICS OF WING-BODY COMBINATIONS 
NEAR THE SPEED OF SOUND. Richard T. 
Whitcomb. September 1952. 41p. diagrs., photos., 
3 tabs. (NACA RM L52H08) (Declassified from 
Confidential, 7/26/55) 

Results are presented which indicate that near the 
speed of sound the zero-lift drag rise of a thin low- 
aspect-ratio wing -body combination is primarily de- 
pendent on the axial distribution of the cross- 
sectional areas normal to the airstream. Results 
of an investigation of applications of this concept to 
the reduction of the drag-rise increments of repre- 
sentative wing-body combinations are also presented. 



NACA RM L54A29a 

ON SLENDER-BODY THEORY AT TRANSONIC 
SPEEDS. Keith C. Harder and E. B. Klunker. 
March 1954. 12p. (NACA RM L54A29a) 
(Declassified from Confidential, 9/7/55) 

The basic ideas of the slender-body approximation 
have been applied to the nonlinear transonic -flow 
equation for the velocity potential in order to obtain 
some of the essential features of slender-body theory 
at transonic speeds. The results of the investigation 
are presented from a unified point of view wliich 
demonstrates the similarity of slender-body solu- 
tions in the various Mach number ranges. The tran- 
sonic area rule and some conditions concerning its 
validity follow from the analysis. 



NACA - Langley Field, Vi.